Biography

My technical background is in aerospace structures and mechanisms, and my activities span from academic research & teaching to consultancy for industry, covering both theoretical work and experimental activity. In addition, the natural progression of my career has increased my management responsibilities, giving me the opportunity to lead multidisciplinary research groups, developing strategies to ensure sustainable growth and supervising their implementation managing the appropriate resources. Currently as Director of the centre my responsibilities cover all its research groups (approximately 90 individuals, including academics, research/support staff and postgraduate researchers), working on a range of topics from space missions design and delivery, to hardware development (from concept design & analysis, to manufacturing assembly and testing) to downstream applications, including our educational & knowledge transfer activities.

Graduated with a first class honour degree in Aerospace Engineering my career started as a stress analyst at Alenia Spazio (Turin I) working at the Columbus module of the International Space Station, then worked at ESA-ESTEC (Noodrwijk NL) on various payloads for the Space Shuttle programme, and in 1995 moved to the University of Southampton where I obtained a PhD and later a Lectureship in Aerospace Structural Dynamics. Progressed though my academic career and I was awarded a personal Chair in Aerospace Structures becoming the head of the Astronautics Research Group.

In 2013 I moved to the Surrey Space Centre, to take up the Royal Academy of Engineering Research Chair in Space Engineering co-sponsored by Surrey Satellite Technology, and later became director of the Centre.

Some of the results of my academic research work have found applications in various industrial hardware developments, from satellites subsystems to whole satellite analysis and testing related activities, contributing to European Space Agency guidelines. My academic work is described in over 200 publications, and has been presented at the major conferences, including invited and plenary key note talks. More details in the Research section.

In parallel with my academic work I have carried out consultancy work on various aspects of spacecraft structures for most of the key players in the UK space industry, SSTL, Astrium (now Airbus defence and Space), MSSL, SSBV, SULA Systems, Lockheed Martin, ABSL, JCR Systems, SIL.

In addition I also lead the development and execution, of space missions to demonstrate technologies ranging from space debris removal (e.g. the RemoveDebirs mission) to in orbit demonstration of hardware developed by UK companies (e.g. AlSat-1N) supporting knowledge transfer towards developing countries.

Research

Research interests

My research work has been sponsored by ESA, TSB; EPSRC, Royal Academy of Engineering, Nuffield Foundation, UK Space Agency and various companies operating in the space sector, such as SSTL and Astrium (now Airbus defence and Space).

Some of the topics covered in this research are: active and passive control of microvibrations, mechanical design of electronic equipment for aerospace applications, multifunctional spacecraft structures; deployable structures; spacecraft stable structures, vibration testing & isolation, actuators design, and FEM validation. Some of the work has directly contributed to ESA flagship missions like BepiColombo, and successful commercial satellites like the SSTL DMC3 constellation. This work is described in over 200 publications, and has been presented at the major conferences, including invited and plenary key note talks. More details of my research work can be found in the Spacecraft Structures Materials and Mechanisms research group web page.

Besides my personal academic research I lead the development and execution of experimental space missions. Two representative examples are:

The RemoveDebirs mission, to demonstrate technologies for space debris removal, that has attracted significant media attention with interviews from major international news outlets:, CNN, BBC, The Times, The Guardian, Wall Street Journal, etc. ).

The Alsat 1N mission, to demonstration in orbit of hardware developed by UK companies, whilst simultaneously acting as a vehicle for a Knowledge Transfer programmes to support the development of the space sector in Algeria.

Test planning and post-test correlation activity have been issues of growing importance in the last few decades and many methodologies have been developed to either quantify or improve the correlation between computational and experimental results. In this article the methodologies established so far are enhanced with the implementation of a recently developed procedure called Virtual Testing. In the context of fixed-base sinusoidal tests (commonly used in the space sector for correlation), there are several factors in the test campaign that affect the behaviour of the satellite and are not normally taken into account when performing analyses: different boundary conditions created by the shaker?s own dynamics, non-perfect control system, signal delays etc. All these factors are the core of the Virtual Testing implementation, which will be thoroughly explained in this article and applied to the specific case of Bepi-Colombo spacecraft tested on the ESA QUAD Shaker. Correlation activity will be performed in the various stages of the process, showing important improvements observed after applying the final complete methodology.

The modal assurance criterion (MAC) and normalized cross-orthogonality (NCO) check are widely used to assess the correlation between the experimentally determined modes and the finite element model (FEM) predictions of mechanical systems. Here, their effectiveness in the correlation of FEM of two types of multi-physics systems, namely, viscoelastic damped systems and a shunted piezoelectric system are investigated using the dynamic characteristics obtained from a nominal FEM, that are considered as the ?true? or experimental characteristics and those obtained from the inaccurate FEMs. The usefulness of the MAC and NCO check in the prediction of the overall loss factor of the viscoelastic damped system, which is an important design tool for such systems, is assessed and it is observed that these correlation methods fail to properly predict the damping characteristics, along with the responses under base excitation. Hence, base force assurance criterion (BFAC) is applied by comparing the ?true? dynamic force at the base and inaccurate FEM predicted force such that the criterion can indicate the possible error in the acceleration and loss factor. The effect of temperature as an uncertainty on the MAC and NCO check is also studied using two viscoelastic systems. The usefulness of MAC for the correlation of a second multi-physics FEM that consists of a shunted piezoelectric damped system is also analyzed under harmonic excitation. It has been observed that MAC has limited use in the correlation and hence, a new correlation method ? current assurance criterion ? based on the electric current is introduced and it is demonstrated that this criterion correlates the dynamic characteristics of the piezoelectric system better than the MAC.

One of the most significant drivers in satellite design is the minimization of mass, in the attempt to reduce the large costs involved in the launch. With technological advances across many fields it is now widely known that very low mass satellites can perform a wide variety of missions. However, the satellite power requirement does not reduce linearly with mass, creating the need for efficient and reliable small satellite deployable structures. One structural solution for this application is tape springs. Tape springs have been previously studied by many countries for space applications focusing on two dimensional systems. This work studies the possible impact of using tape springs folded in three dimensions. By initially analytically determining the static moments created, simple deployment models can be constructed for tape springs in free space. By determining the impact of these moments about an array fold line, a dynamic model of an array can be created which is directly comparable to the two dimensional system. The impact of the three dimensional fold can then be determined.

One of the most significant drivers in satellite design is the minimization of mass to reduce the large costs involved in the launch. With technological advances across many fields, it is now widely known that very low-mass satellites can perform a wide variety of missions. However, there is a need for small, efficient, area deployment devices. One possible structural solution for such devices is tape springs. Previous work on tape spring hinges has focused on two-dimensional folds; however, applications exist that incorporate three-dimensional tape spring folds. The properties of three-dimensional tape spring folds are experimentally investigated using a specially designed test rig. The rig is first used to produce comparative two-dimensional data before being used to analyze more complex three-dimensional folds.

The ROV-E project is a three year European Union Framework 7 project, which began in January 2011, dedicated to the research and development of lightweight technologies for exploration rovers. As part of this the University of Southampton, along with other consortium members, have been looking into the development of a Multifunctional Power Structure (MFPS). This is a structure that combines aspects of the electrical power system into a single panel component, removing the unnecessary mass of additional structures and containers required to support distributed discrete components inside a rover. The specific components imbedded into the multifunctional panel include: power generation (photovoltaic cells), control electronics and power storage. The main focus of the research at the University of Southampton was the power storage function of the panel, which aimed at exploiting the cost benefits of using off the shelf components by using commercially available lithium polymer battery cells. Initial validation testing exposed these cells to structural, temperature and pressure environments which proved the robustness of the cells throughout the predicted lifecycle of the multifunctional panel. An initial representative honeycomb panel incorporating battery cells was constructed to validate the manufacturing process. This panel was then used experimentally to assess the failure methods of the cells, revealing that the cells are more likely to suffer performance loss due to bending than accelerations. Following on from the initial validation testing a full MFPS was designed and optimised before being subjected to mechanical and thermal environments. This paper focuses on the final design and testing of this complete MFPS. Although the testing encountered various unforeseen problems, the batteries were both mechanically and thermally validated as part of the complete MFPS.

Microvibrations, generally defined as low amplitude vibrations at frequencies up to 1 kHz, are now of critical importance in a number of areas. One such area is on-board spacecraft carrying sensitive payloads, such as accurately targeted optical instruments or micro-gravity experiments, where the microvibrations are caused by the operation of other equipment, such as reaction wheels, necessary for its correct functioning. It is now well known that the suppression of such microvibrations to acceptable levels requires the use of active control techniques which, in turn, require sufficiently accurate and tractable models of the underlying dynamics on which to base controller design and initial performance evaluation. Previous work has developed a systematic procedure for obtaining a finite-dimensional state space model approximation of the underlying dynamics from the defining equations of motion which has then been shown to be a suitable basis for robust controller design. This modeling approach is based on the use of Lagrange's equations of motion and is one of a number of possible models possible in this area. In this paper, we describe the experimental validation of this model prior to experimental studies to determine the effectiveness of the designed controllers with the objective of establishing the effectiveness of this procedure both stand alone and against alternatives.

This paper reflects on some of the experiences of correlating the Finite Element Model (FEM) of ESA's BepiColombo spacecraft to sine sweep test measured data, in order to validate the model for subsequent Coupled Loads Analysis (CLA). Post-model update procedures can take a significant number of man hours to complete, without necessarily resulting in a final FEM which is notably more representative of the real structure than the initial FEM. The long term research intention is to use the lessons learnt from BepiColombo and other spacecraft correlations to work towards the containment of the FEM correlation process, this paper addresses part of this on-going research effort. This is to be achieved through: investigating the current techniques/algorithms for FEM test analysis correlation and validation; identifying their limitations; thus ultimately developing a methodology which identifies where model updates are or are not necessary to achieve a model which is validated for key participating modes in flight predictions or during any sine test notching.

Microvibrations, at frequencies between 1 and 1000 Hz, generated by on board equipment, propagate throughout a spacecraft structure affecting the performance of sensitive payloads. The purpose of this work is to investigate strategies to model and reduce these dynamic disturbances by active control. Initial studies were performed by considering a mass loaded panel where the disturbance excitation source consisted of point forces, the objective being to minimise the displacement at an arbitrary output location. Piezoelectric patches acting as sensors and actuators were used. The equations of motion are derived by using Lagrange's equation with modal shapes as Ritz functions. The number of sensors/actuators and their location is variable. The set of equations obtained is then transformed into state variables and some initial controller design studies have been undertaken. These are based on feedback control implemented using a full state feedback and an observer which reconstructs the state vector from the available sensor signal. Here, the basics behind the structural modelling and controller design will be described. This preliminary analysis will also be used to identify short to medium term further work.

The results of an extensive test program to characterize the behavior of typical aircraft structures under acoustic loading and to establish their fatigue endurance are presented. The structures tested were the three flap-like box-type of structures. Each structure consisted of one flat (bottom) and one curved (top) stiffener stiffened skin panel, front, and rear spars, and ribs that divided the structures into three bays. The three structures, constructed from three different materials (aircraft standard aluminum alloy, Carbon Fibre Reinforced Plastic, and a Glass Fibre Metal Laminate, i.e., GLARE) had the same size and configuration, with only minor differences due to the use of different materials. A first set of acoustic tests with excitations of intensity ranging from 140 to 160 dB were carried out to obtain detailed data on the dynamic response of the three structures. The FE analysis of the structures is also briefly described and the results compared with the experimental data. The fatigue endurance of the structures was then determined using random acoustic excitation with an overall sound pressure level of 161 dB, and details of crack propagation are reported.

Multifunctional structures offer savings in spacecraft mass and volume by combining the functional elements of subsystems with structural components. However, the benefit of these mass and volume savings can be outweighed by the cost of the materials and processes required for their manufacture. This paper presents a compromise, taking advantage of some of the mass and volume savings associated with multifunctional structures, whilst remaining relatively low in price. By taking commercial electrical cells, and integrating them into the structure, the parasitic mass of the power subsystem is eliminated and the structural properties of the cells harnessed to produce a multifunctional powerstructure. In this work, prismatic plastic lithium-ion cells are used as part of the core of a carbon fibre sandwich panel, along with aluminium honeycomb. This makes use of the batteries' structural properties and also removes their volume from the spacecraft bus to an area that would otherwise be filled with inert material. The paper assesses the feasibility of this concept. Firstly, a vibration test has been successfully conducted to prove the ability of such cells to survive the launching environment. Secondly, a panel has been successfully manufactured using these cells as described, without serious damage to the cells' performance.

Foster JA, Aglietti GS(2012)Strategies for Thermal Control of a Multifunctional Power Structure Solar Array, Journal of Aerospace Engineering25(3) ASCE

This paper develops a modelling technique for equipment load panels which directly produces (adequate) models of the underlying dynamics on which to base robust controller design/evaluations. This technique is based on the use of the Lagrange's equations of motion and the resulting models are verified against those produced by a finite Element Method Model.

This paper illustrates a procedure to calculate the response of a tethered spherical aerostat to gusts, including the effect of structural nonlinearity and accounting for some of the fluid - structure interaction between the aerostat and tether line. The procedure developed and presented here is based on a full three-dimensional dynamic finite element model, with aerodynamic loads calculated from the relative velocity between a time-varying input airflow and resulting structural velocities. Exact solutions for the static response and a simplified dynamic model, both developed to validate the results of the procedure illustrated in this paper, are also derived and described in detail. The dynamic responses to gusts are compared with the equivalent steady-state solution to assess the approximations of the static solutions. Particular emphasis is placed on the output rotation of the aerostats to quantify disturbances on the pointing stability produced by gusts.

Deployable structures are required for many satellite operations, to deploy booms for communications or area deployment for power generation, and many sophisticated mechanisms have been developed for these types of structures. However, tape springs, defined as thin metallic strips with an initially curved cross-section, are an attractive structural solution and hinge mechanism for satellite deployable structures because of their low mass, low cost and general simplicity. They have previously been used to deploy booms and array panels in various configurations that incorporate small two-dimensional tape hinges, but they also have the potential to be used in greater numbers to create larger, more geometrically complicated deployable structures. The aim of this work is to investigate a computationally efficient method of simulating these tape spring based deployable structures and to determine the limitations of the analysis approach. The study focuses on a specific deployable structure layout that incorporates 148 separate tape spring elements in three fold lines. The static and dynamic properties of the system are initially investigated experimentally allowing the basic parameters of the theoretical model to be determined accurately. It was found that the simulated tape pair stiffness was a key parameter affecting the dynamic properties of the model and the peak shock accelerations. It was concluded that the model was capable of closely simulating the dynamic 'snap through' behaviour of the wall. However, the torsional stiffness around the axis normal to the plane of the structure was found to be too large, resulting in over approximations of the peak shock accelerations.

Inflatable technology for space applications is under continual development and advances in high strength fibres and rigidizable materials have pushed the limitations of these structures. This has lead to their application in deploying large-aperture antennas, reflectors and solar sails. However, many significant advantages can be achieved by combining inflatable structures with structural stiffeners such as tape springs. These advantages include control of the deployment path of the structure while it is inflating (a past weakness of inflatable structure designs), an increased stiffness of the structure once deployed and a reduction in the required inflation volume. Such structures have been previously constructed at the Jet Propulsion Laboratory focusing on large scale booms. However, due to the high efficiency of these designs they are also appealing to small satellite systems. This article outlines ongoing research work performed at the University of Southampton into the field of small satellite hybrid inflatable structures. Inflatable booms have been constructed and combined with tape spring reinforcements to create simple hybrid structures. These structures have been subjected to bending tests and compared directly to an equivalent inflatable tube without tape spring reinforcement. This enables the stiffness benefits to be determined with respect to the added mass of the tape springs. The paper presents these results, which leads to an initial performance assessment of these structures.

Contracted by the European Commission in the frame of the EU?s Seventh Framework
Programme for Research (FP7), a wide European consortium has been working since 2013
towards the design of a low cost in-orbit demonstration called RemoveDEBRIS. With a targeted
launch date in the second quarter of 2016, the RemoveDEBRIS mission aims at demonstrating
key Active Debris Removal (ADR) technologies, including capture means (net and harpoon
firing on a distant target), relative navigation techniques (vision-based navigation sensors and
associated algorithms), and deorbiting technologies (drag sail deployment after the mission
followed by an uncontrolled reentry). In order to achieve these objectives, a micro satellite testbed
will be launched into a Low Earth Orbit, where it will deploy its own dedicated targets and
CubeSats to complete each demonstration. As part of its System Engineering role, Airbus
Defence and Space has been conducting the Mission Analysis studies for this unprecedented
mission. This paper will present a description of the RemoveDEBRIS demonstration objectives
and scenario and will present in detail some specific mission related analyses and trade-offs that
have driven the mission design.

This paper investigates the technical feasibility of a system that could be used to collect the solar irradiation at high altitude, convert it into electricity, and then transmit it to the ground via a cable. As a first step to assess the viability of this device, an estimate of the solar irradiation that can be expected at a defined altitude above the ground is presented, based on real atmospheric data. The study demonstrates that locating PV devices at high altitude with the use of an aerostatic platform, could bring a significant advantage in the production of electrical power, if compared with a typical UK ground based PV system. The fundamental equations for a preliminary design of the system are presented together with a first realistic choice of the most relevant engineering parameters that need to be set. An estimate of the cost of the system is provided and the possible risks involved, applications, advantages and disadvantages of the technology are assessed.

The issue of model reduction is one that must often be overcome in order to perform the necessary checks as part of the spacecraft Finite Element Model (FEM) validation process. This work compares different reduction methods; specifically the popular and long-standing Guyan method, and the potentially more accurate System Equivalent Reduction Expansion Process (SEREP). The influence of sensor set location on the quality of the reduced model has also been considered, and the commonly applied methods to maximize kinetic energy and effective independence have been applied. These investigations have taken the form of studies involving two large, unique, scientific spacecraft. The computational results are compared with experimental results that are also detailed in the paper. The findings highlight the potential issues with the accuracy of a Guyan reduced model in replicating the full system dynamics, even with a reasonably large sensor set. It is shown that this can be improved slightly in some circumstances through implementation of sensor set placement optimization techniques. The SEREP method is shown to have the benefit of being more accurate at replicating the full system behavior than the more traditional Guyan method, while also producing higher diagonal values in cross-orthogonality comparisons between FEM and test.

In this study, the effectiveness of model correlation methods such as modal assurance criterion (MAC) and normalized cross orthogonality check (NCO) in the prediction of forced response characteristics of spacecraft structure are assessed using synthetic modal parameters as well as physically altered finite element models of two real spacecraft structures. It is observed that, neither MAC nor NCO can assure the finite element model capability to represent the acceleration response on the structure and the force transmitted to the base within the acceptable limits during the base excitation. It is also observed that, sometimes model with lower values of MAC or NCO represent the response characteristics better than that of a model with high value of MAC or NCO. A synthetic model is used for the robustness study of spacecraft models under base excitation and this model can approximately represent the response characteristics.

In this paper attention is focused on a simply supported panel, with twin patches of piezoelectric material bonded on opposite faces of the panel acting as actuators. A test rig comprising of an aluminium alloy panel has been designed and built. Particular attention has been placed in designing the rig to reproduce as accurately as possible a simple support along all four edges. The deign and analysis of the rig were carried out using the Finite Element (FE) method, and the results of the FE analysis are then compared and validated against the experimental results. A Mechanical Impedance based method, and the Lagrange Rayleigh-Ritz Method were then used to produce mathematical models of the actively controlled panel. These two techniques are chosen as representative of commonly used techniques in the production of mathematical models for active control design studies. The results obtained from the numerical simulations were compared with experimental results in order to assess the accuracy and sensitivity of the modelling techniques.

Schwingshackl CW, Cunningham PR, Aglietti GS(2004)Honeycomb elastic material properties: A review of some existing theories and a new dynamic approach, Proceedings of the 2004 International Conference on Noise and Vibration Engineering, ISMApp. 1353-1363

The influence of the nine orthotropic material properties of honeycomb on the dynamic response of a finite element model of a simple supported sandwich plate are examined. Fifteen available theories from the literature for the material properties of honeycomb are reviewed and their values calculated for a Hex Web 5.2-1/4-25(3003) Aluminium core. The agreement between the theoretical material properties and the major ASTM (American Society for Testing and Materials) standard test methods is investigated. A new and simple technique is described for measuring the dynamic shear moduli of honeycomb materials and its values are compared with those presented in the literature.

Purpose - The purpose of this paper is to assess the suitability of various methods for the reduction of a large finite element model (FEM) of satellites to produce models to be used for correlation of the FEM with test results. The robustness of the cross-orthogonality checks (COC) for the correlation process carried out utilizing the reduced model is investigated, showing its dependence on the number of mode shapes used in the reduction process. Finally the paper investigates the improvement in the robustness of the COC that can be achieved utilizing optimality criteria for the selection of the degrees of freedom (DOF) used for the correlation process. Design/methodology/approach - A Monte Carlo approach has been used to simulate inaccuracies in the mode shapes (analysis and experimental) of a satellite FEM that are compared during the COC. The sensitivity of the COC to the parameters utilized during the reduction process, i.e. mode shapes and DOFs, is then assessed for different levels of inaccuracy in the mode shapes. Findings - The System Equivalent Expansion Reduction Process (SEREP) has been identified as a particularly suitable method, with the advantage that a SEREP reduced model has the same eigenvalues and eigenvector of the whole system therefore automatically meeting the criteria on the quality of the reduced model. The inclusion of a high number of mode shapes in the reduction process makes the check very sensitive to minor experimental or modelling inaccuracies. Finally it was shown that utilizing optimality criteria in the selection of the DOFs to carry out the correlation can significantly improve the probability of meeting the COC criteria. Research limitations/implications - This work is based on the FEM of the satellite IT>Aeolus/IT>, and therefore the numerical values obtained in this study are specific for this application. However, this model represents a typical satellite FEM and therefore the trends identified in this work are expected to be generally valid for this type of structure. Practical implications - The correlation of satellite FEM with test results involves a substantial effort, and it is crucial to avoid failures of the COC due to numerical issues rather than real model inaccuracies. This work shows also how an inappropriate choice of reduction parameters can lead to failure of the COC in cases when there are only very minor differences (e.g. due to minor amount of noise in the results) between analytical and test resul

A self-contained inflatable and rigidizable truss based substructure, its constraining mechanism, and stowage enclosure were developed for the RemoveDEBRIS technology demonstrator. RemoveDebris is a European Commission FP7 funded mission due for launch in late 2016. The hardware discussed in this paper will be integrated with the DebrisSat-1 microsatellite. During the course of the mission, active debris removal will be achieved by capturing DebrisSat-1 with the aid of a net fired from the primary platform. The inflatable module is key to this experiment as it allows the simulation of a much larger piece of debris than would be possible with a CubeSat alone. Following its capture, the inflatable structure will continue with its second objective as an end of life removal solution by passively drag augmenting DebrisSat-1's orbit to re-entry. The inflatable structure is constructed with six aluminum-polymer laminate cylindrical booms. These are connected in an axial manner to form a regular octahedron with a cross sectional area of 0.5 m2. A set of eight triangular polyester film segments or sails enclose the structure. The segments serve a dual purpose: firstly to increase the aerodynamic drag of the spacecraft, and secondly to distribute impact loads between the compressive inflatable members. A single cool gas generator (CGG) is utilised to deploy and rigidize the structure. This paper examines the development of the inflatable module from the early conceptual stages to the pre-qualification test level.

This paper addresses the problem of actively attenuating the vibration of plates on satellites. A pure feedback controller is demonstrated which operates at a set of dominant frequencies in a disturbance spectrum, where the control path model is estimated on-line. A new feature of the frequency selective feedback is the use of the inverse Hessian to improve adaptation speed. The control scheme also incorporates a frequency estimation technique to determine the relevant disturbance frequencies with higher precision than the standard fast Fourier transform (FFT). The controller is implemented on a test rig to demonstrate the practical feasibility of the method. A disturbance with three rational dominant frequencies is introduced. If FFT were used instead of the frequency estimation method proposed, then a large number of samples would be required to accurately estimate the disturbance frequencies, and, most importantly, FFT-estimated frequencies could lead to an unstable control system due to their granularity. Using the proposed frequency estimation method, the total achieved attenuation is 26dB on the experimental rig.

Accurate resolution imaging from Satellites involves large amounts of data that has to be stored on board the spacecraft computer. The data can be stored on Hard Disc Drives. However survival to the mechanical environment existing during the spacecraft launch and to the space environment during satellite operations are two major challenges in the use of HDDs for spacecraft applications. This paper describes the process that generated the design of an enclosure that has allowed conventional Personal Computer's HDDs to be used on board the SSTL spacecraft BEIJING-1. The design philosophy is discussed and the extensive test campaigns that supported the selection of a suitable HDD are described. The focus of the work was the design and implementation of a suspension system to reduce to acceptable levels the random vibration environment experienced by the HDD. The tests carried out on the suspension system showed that this was able to reduce by approximately 50% the rms acceleration experienced by the units. Thus allowing their use on the spacecraft. The spacecraft was launched in October 2005, and to date the HDD units are operating correctly.

Meurers T, Veres SM, Aglietti G, Rogers E(2003)Frequency selective microvibration control for a single plate, Proceedings of the Tenth International Congress on Sound and Vibrationpp. 179-186

In recent years, frequency selective controllers were developed to offer an alternative to the well studied and widely used filtered reference least mean square algorithm. The motivation came from the fact, that many disturbance spectra are dominated by a finite set of frequencies (tones). Hence it was concluded that noticeable reduction could be achieved by just controlling the dominant tones. The main advantages of frequency selective controllers are that stability can be shown easily and the controller is of lower order compared to standard FIR filters. Initially, the here presented controllers were tested in sound experiments and excellent behaviour was demonstrated. Then a vibrating plate test rig was acquired and it was possible to move on to examine the control behaviour for microvibrations. A lump mass was fixed to the plate to create an unbalanced behaviour. Furthermore, 4 piezoelectric components were glued to the plate. The arrangement is, that 2 piezoelectric elements are opposite each other on either side of the plate. Different experimental trials were carried out. It was started with driving one piezoelectric element as a disturbance generator and the use of two more as actuator and accelerometer respectively. Model free and model based controllers were implemented. All controller types achieved an overall attenuation for the whole frequency band of more than 10 dB, not only for the controlled frequencies justifying the approach taken.

Micro-vibrations on board spacecraft are an issue of growing importance, as some modern payloads, and in particular the new generations of optical instruments require extreme platform stability. These low level mechanical disturbances are usually created by the functioning of mechanical equipment (sources) such as reaction wheels, antenna pointing mechanisms cryo-coolers etc., and cover a wide frequency range. Because of the low level of the vibrations and their wide frequency range, the modeling and analysis of micro-vibrations poses a challenge as the typical structural modeling techniques used in this sector (Finite Element Method (FEM) and Statistical Energy Analysis (SEA)) are reliable only in some areas of the frequency spectrum. The FEM is well suited for low level frequencies; whereas energy methods (e.g. SEA or Energy Finite Element Method EFEA) are suited for high-frequency problems; in the mid-frequency range, finally, other methods (e.g. Hybrid FEA-SEA) tend to be used, even if they're still not well-established such as the ones named before. However the issue is that there is no single method that can address micro-vibrations in the whole frequency range. In this paper, the methods cited above will be very briefly reviewed and their use in specific micro-vibration prediction problems will be investigated in detail and compared with experimental results. In practice the work presented here uses the Finite Element Method as base-line method to investigate the whole frequency range (say up to 1000 Hz). The FEM predictions are then compared with the experimental results, showing that at medium and high frequencies the response start to deviate significantly from the FEA predictions. The high frequency behavior of the structure will be analyzed using SEA. The mid-frequency range, finally, will be tackled from both directions: from the high frequency side using the Hybrid FE-SEA, whereas from the low frequency side the capability of the standard FEM will be extended using stochastic FEM. The tests are carried out using the structural qualification model of an SSTL satellite bus that has been used to support a high resolution camera. The computational transfer functions and those from the experimental activity will be finally compared using the Modal Assurance Criteria (MAC).

In this paper, the conventional design of an enclosure for electronic equipment for space application is reviewed and an alternative type of construction is proposed. The alternative design is based on the use of carbon fibre reinforced plastic (CFRP) sandwich panels for the construction of the enclosure, and the substitution of the printed circuit board (PCB) antivibration frames (AVFs) with antivibration rods (AVRs). To put this work into context, the requirements applicable to the structural design of this type of unit are briefly reviewed. Standard structural analyses have been performed on the conventional enclosure and then repeated for the proposed configuration, in order to demonstrate its compliance with the fundamental mechanical requirements. The issues concerning the radiation protection offered by the enclosure are discussed, and some solutions to this potential problem are briefly presented. The work demonstrates the possibility of achieving a saving of about 20 per cent on the overall mass of the unit. Finally, the cost of the proposed enclosure is assessed and compared with the conventional design for various missions.

Dynamic models are widely used in many branches of science and engineering, and it has been argued that many of the shortfalls with these models are due to the fact that the physics of joint dynamics are not fully understood. This makes the phenomenon very hard to model theoretically from first principles. Experimental analyses are therefore widely used to underpin any work in this area. This study aims to build on the previous experimental work based on simple beam joints and analyzes the damping trends for metal panels. This incorporates torsional effects into the system creating more complex displacements of the joint. Five panel configurations are investigated using an experimental approach that minimizes all external influences on the dynamics of the panels. Each mode loss factor is determined from these experimental tests and compared with the most established theoretical model in the field. The corresponding joint displacements and decay trends are also analyzed, producing indications as to the likely dominant source of damping, suggesting that the mode shapes can be categorized based on their displacement and dominant damping source.

An effective investigation of alternative control strategies for the reduction of vibration levels in satellite structures requires realistic, yet efficient, structural models to simulate the dynamics of the system. These models should include the effects of the sources, receivers, supporting structure, sensors, and actuators. In this paper, a modeling technique which meets these requirements is developed and some active control strategies are briefly investigated. The particular subject of investigation is an equipment-loaded panel and the equations of motion are derived using the Lagrange-Rayleigh-Ritz (LRR) approach. The various pieces of equipment on the panel are mounted on active or passive suspensions, and resonators are used to represent the internal dynamics of the mounted equipment. Control of the panel, which transmits vibrations from sources to receivers, is by means of piezoelectric patches and the excitation consists of dynamic loads acting on the equipment enclosures and/or directly on the panel. The control objective is to minimize the displacement at an arbitrary output location. The LRR model developed is verified against one produced by using the finite-element method. Finally, some initial controller design studies are undertaken to investigate and compare the effectiveness of different control strategies (e.g., minimization at the source, along the vibration path, or at the receiver).

This paper describes the validation of a Lagrange-Rayleigh-Ritz technique for the mathematical modeling of mass loaded panels for active control studies. The validation has been carried out by comparing the results produced by a Lagrange-Rayleigh-Ritz model with the results produced by a finite-element model and experimental data. Attention was focussed on a simply supported panel with a lumped mass constrained to its surface to simulate the presence of equipment mounted on the panel and twin piezoelectric patches bonded to the panel, working as sensors and actuators. The design of the experimental rig is described in detail, and a test campaign was carried out to obtain a set of transfer functions characteristic of this plant. The experimental data are then used to validate the predictions of the mathematical model. In particular, it was demonstrated that the Lagrange-Rayleigh-Ritz model was able to reproduce accurately the dynamics of the plant requiring a relatively small number of degrees of freedom.

Highly stable structures that are destined for space use are vulnerable to dimen-sional stability loss due to random vibration loads experienced during launch and ground testing. Small movements at structural interfaces and non-recoverable strains induced in me-tering elements can have negative implications for optical performance on-orbit. Often the dimensional stability aspects of optical bench structures are verified by environ-mental tests on Engineering or Protoflight Model instruments. It is proposed that a better understanding of vibration-induced structural dimensional stability loss could enable the as-sessment of stability loss through analysis at an early stage. To this end, several tests have been developed at RAL to assess dimensional stability loss in materials and joints in a controlled manner under random vibration. One test has been used to assess Al alloy and CFRP material samples in a 4-point bending configuration, and anoth-er has been used to assess micron-level slipping at a bolted interface. The aim of these tests was to provide useful material data, and also to assess the feasibility of predicting stability loss caused by random vibration events. It was found that the classical frequency domain random vibration Finite Element Analysis commonly performed to assess safety margins against structural failure on space structures is probably insufficient to predict instability. This is because the results are highly dependent on non-symmetry in the stress response (ie, due to gravity, pre-stress, or non-linear response). However a good correlation with test results was achieved with a time domain FEA model which incorporates nonlinear kinematic hardening rules in the materials.

Micro-vibration is a low level disturbance, which cannot be controlled or reduced by the Attitude and Orbit Control System of a spacecraft. It can emanate from various sources on a typical spacecraft, notably subsystems with moving parts such as reaction wheels or cooler mechanisms. Micro-vibration can also result from thermo-elastic effects due to stick-slip from differential expansion of parts. It causes problems for sensitive payloads especially high resolution cameras where the demand for higher resolution (which drives stability requirements), has made analysis and control of microvibrations relevant for a larger number of satellites. The availability of mathematical models to represent the disturbance sources in a format that can themselves be coupled to the mathematical model of the structure to perform end-to-end analysis to obtain predictions of the stability level at the receiver is crucial. In the simplest form the sources can be represented with forces and moments of appropriate characteristics, adding also some inertia at the source location. But even for this simplistic implementation it is necessary to have available the details of the forces and moment produced by a source, and these can be either calculated from schematic models of the functioning of the device, or experimentally determined, or a mixture of the two. Typical test techniques applied to micro-vibration measurement /characterization will be described, highlighting advantages and drawbacks of the various methods. A simple experimental apparatus for the measurement of the micro-vibrations emitted by a reaction wheel is presented. A mathematical model of the reaction wheel disturbances is also presented together with its coupling to a typical spacecraft structural model.

The emerging field of multifunctional structure (MFS) technologies enables the design of systems with reduced mass and volume, thereby improving their overall efficiency. It requires developments in different engineering disciplines and their integration into a single system without degrading their individual performances. MFS is particularly suitable for aerospace applications where mass and volume are critical to the cost of the mission. This article reviews the current state of the art of multifunctional structure technologies relevant to aerospace applications.

One of the most significant drivers in satellite design is the minimization of mass to reduce the large costs involved in the launch. With technological advances across many fields, it is now widely known that very low-mass satellites can perform a wide variety of missions. However, the satellite power requirement does not reduce linearly with mass, creating the need for efficient and reliable small satellite deployable structures. One possible structural solution for this application is tape springs. Tape springs have been previously studied in many countries for space applications focusing on two-dimensional systems. This work studies the possible impact of using tape springs folded in three dimensions. By first analytically determining the static moments created, simple deployment models can be constructed for tape springs in free space. Then, determining the impact of these moments about an array fold line allows the creation of a dynamic model of an array that is directly comparable to the two-dimensional system. The impact of the three-dimensional fold can then be determined.

A main aim of spacecraft design is cost reduction that can either be achieved by reducing the cost of the spacecraft or by lowering the cost to launch. One proposed technology to reduce the mass and therefore lower the launch costs is the multifunctional power structure concept that incorporates the secondary spacecraft power supply into the load carrying structure. Here a short introduction of the dynamic analysis and optimization of such a structure is presented. The manufacture and testing of a multifunctional power panel is discussed in detail and its dynamic response is compared to a conventional honeycomb panel. The multifunctional design successfully combined the structural and power storage functions. It provided a similar dynamic response to conventional spacecraft structures and improved the energy density.

In the current world of engineering, structural vibration problems continue impact the design and construction of a wide range of products. Amid the parameters that determine the dynamic behaviour of a structure the one that takes into account the dissipation of energy resulting in the decay of the vibration is the least understood and the most difficult to quantify [1]. The estimation of damping factors is of interest in most branches of engineering sciences. In the field of aircraft structures the damping directly affects the fatigue life, a parameter which is applied conservatively due to the inherent complexity in modelling the damping of built up structures and the potentially catastrophic consequences of a fatigue failure. One of the most important problems is the limited knowledge of how joints affect the damping of the complete structure. This work therefore addresses this issue and focuses on the damping of joints in metal plates as part of a larger project to investigate the damping of built up structures. Various plate configurations are experimentally investigated using two different approaches. The results from the configurations are compared and discussed along with the advantages and disadvantages of each experimental approach. This enables a link to be identified between the damping magnitudes and the mode shapes and joint stiffnesses.

Currently, vibroacoustic problems can be solved using a wide range of numerical techniques. In the low-frequency range, element-based deterministic methods, such as the Finite Element Method (FEM) and Boundary Element Method (BEM) are regularly employed to define the structural and acoustic domains, respectively. The fully coupled FEM-BEM is a classic, vastly popular method. In the high-frequency range probabilistic methods, such as Statistical Energy Analysis, tend to be more efficient and produce more reliable results. Although new techniques are becoming available (e.g. Hybrid FE-SEA Method), the characterisation of the mid-frequency behaviour still poses some challenges, as the computational cost of element-based techniques is often prohibitive, and the modal density is not sufficiently high for statistical approaches to be applicable.
This paper discusses an approach aimed at improving the efficiency of the classic FEM-BEM method and potentially extending its usability to the mid-frequency band, specifically in the context of space-craft structural design. The iterative coupling between Craig-Bampton reduced finite element models and BEM is considered as an alternative to directly solving the FEM-BEM coupled equation, allowing the use of efficient procedures for either domain separately. A pre-process enabling the method?s computational implementation is presented, which is based on a manipulation of the reduced mass and stiffness matrices. It is used to allow the application of a distributed load to a Craig-Bampton condensed structure, while mitigating the need to retain a large number of physical degrees of free-dom.
The efficiency of the aforementioned matrix modification procedure is compared to that of performing a full Craig-Bampton reduction, and its cost is expressed in terms of floating point operations. An iterative coupling scheme is used on a test-case structure for both a full physical model and a reduced one to verify the concept, and check whether convergence is susceptible to initial conditions, such as the shape of the acoustic field. Finally, the perturbation of the condensed matrices is shown to produce results consistent with those for the full physical model, while substantially reducing the computational effort required for the simulation.

Deployable structures are required for many satellite operations, to deploy booms for communications or area deployment for power generation, and many sophisticated mechanisms have been developed for these types of structures. However, tape springs, defined as thin metallic strips with an initially curved cross-section, are an attractive structural solution and hinge mechanism for satellite deployable structures because of their low mass, low cost and general simplicity. They have previously been used to deploy booms and array panels in various configurations that incorporate small two-dimensional tape hinges, but they also have the potential to be used in greater numbers to create larger, more geometrically complicated deployable structures. This publication investigates the applicability of using a simplified modelling approach to predict the deployment dynamics of a three dimensional deployable structure that uses a significant quantity of tape springs. This work builds on previous studies which have focused on the analysis of two dimensional tape spring based structures. The configuration being investigated consists of four walls mounted as a square. Each wall has three fold lines allowing the structure to fold down in a concertina style and each fold line is populated by a series of tape spring hinges mounted in pairs. A total number of around 600 individual tape springs elements are used across the 12 fold lines. A computationally efficient method of simulating the three dimensional deployable structure was studied based on a finite element explicit analysis. Equivalent static and dynamic experimental testing on a breadboard structure is presented allowing a direct comparison of the theoretical and experimental data. It was concluded that this simplified analysis approach is capable of modelling the structural dynamics in the deployment direction for three dimensional structural deployments. As a result, the use of this approach could significantly reduce computation time when performing initial design trade-offs.

This article discusses the microvibration analysis of a cantilever configured reaction wheel assembly. Disturbances induced by the reaction wheel assembly were measured using a previously designed platform. Modelling strategies for the effect of damping are presented. Sine-sweep tests are performed and a method is developed to model harmonic excitations based on the corresponding test results. The often ignored broadband noise is modelled by removing spikes identified in the raw signal including a method of identifying spikes from energy variation and band-stop filter design. The validation of the reaction wheel disturbance model with full excitations (harmonics and broadband noise) is presented and flaws due to missing broadband noise in conventional reaction wheel assembly microvibration analysis are discussed.

Active control techniques are often required to mitigate the micro-vibration
environment existing on board spacecraft. However, reliability issues and high
power consumption are major drawbacks of active isolation systems that have
limited their use for space applications. In the present study, an electromagnetic
shunt damper (EMSD) connected to a negative-resistance circuit is designed,
modelled and analysed. The negative resistance produces an overall reduction
of the circuit resistance that results in an increase of the induced current in
the closed circuit and thus the damping performance. This damper can be
classified as a semi-active damper since the shunt does not require any control
algorithm to operate. Additionally, the proposed EMSD is characterised by low
required power, simplified electronics and small device mass, allowing it to be
comfortably integrated on a satellite. This work demonstrates, both analytically
and experimentally, that this technology is capable of effectively isolating typical
satellite micro-vibration sources over the whole temperature range of interest

Aglietti GS, Langley RS, Rogers E, Gabriel SB(2004)Model building and verification for active control of microvibrations with probabilistic assessment of the effects of uncertainties, Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science218(4)pp. 389-400

Microvibrations of a satellite reaction wheel assembly are commonly analysed in either hard-mounted or coupled boundary conditions, though coupled wheel-to-structure disturbance models are more representative of the real environment in which the wheel operates. This article investigates the coupled microvibration dynamics of a cantilever configured reaction wheel assembly mounted on either a stiff or flexible platform. Here a method is presented to cope with modern project necessities: (i) need of a model which gives accurate estimates covering a wide frequency range; (ii) reduce the personnel and time costs derived from the test campaign, (iii) reduce the computational effort without affecting the quality of the results. The method involves measurements of the disturbances induced by the reaction wheel assembly in a hard-mounted configuration and of the frequency and speed dependent dynamic mass of the reaction wheel. In addition, it corrects the approximation due to missing speed dependent dynamic mass in conventional reaction wheel assembly microvibration analysis. The former was evaluated experimentally using a previously designed and validated platform. The latter, on the other hand, was estimated analytically using a finite element model of the wheel assembly. Finally, the validation of the coupled wheel-structure disturbance model is presented, giving indication of the level of accuracy that can be achieved with this type of analyses.

The concept for a Small Satellite Transfer Vehicle is presented. This vehicle is targeted at the deployment of complex groups, constellations or formations of nano- or picosats into an orbit as a secondary payload. The concept provides the capability to deploy this secondary payload into a significantly different orbit from the primary payload on the launch, and has the capability to provide additional services such as spacecraft dispersion along an orbit. It aims to minimise the overheads associated with multiple small spacecraft that are manifested on a single launcher, and offers a number of benefits over the more typical ad-hoc launcher accommodation. The Cubesat PPOD accommodation concept is employed as it provides a suitable open standard that is well used, tested and documented. A mission scenario is introduced and analysed and the system design is described.

Sandwich panels have a very high stiffness to weight ratio, which makes them particularly useful in the aerospace industry where carbon fibre reinforced plastics and lightweight honeycomb cores are being used in the construction of floor panels, fairings and intake barrel panels. In the latter case, the geometry of the panels can be considered doubly curved. This paper presents an introduction to an ongoing study investigating the dynamic response prediction of acoustically excited composite sandwich panels which have double curvature. The final objective is to assess and hopefully produce an up to date set of acoustic fatigue design guidelines for this type of structure. The free vibration of doubly curved composite honeycomb sandwich panels is investigated here, both experimentally and theoretically, the latter using a commercially available finite element package. The design and manufacture of three test panels is covered before presenting experimental results for the natural frequencies of vibration with freely supported boundary conditions. Once validated against the experimental results, the theoretical investigation is extended to study the effects of changing radii of curvature, orthotropic properties of the core, and ply orientation on the natural frequencies of vibration of rectangular panels with various boundary conditions. The results from the parameter studies show curve veering, particularly when studying the effect of changing radii and ply orientation, however, it is not clear whether this phenomenon is due to the approximation method used or occurs in the physical system.

New design approaches will be required to increase the payload to mass fraction for future satellite generations. The multifunctional design concept, where spacecraft subsystems are integrated into the load bearing structure of the satellite, is one considered technology. This paper describes the design, analysis and manufacture of a particular multifunctional power structure with a special focus on its dynamic response. An analytical and a finite element analysis of ten proposed multifunctional power structures, based on a sandwich panel configuration, are presented. The theoretical out-of-plane material properties for the investigated designs are derived with the help of the virtual displacement method. These theoretical properties are compared to finite element models and subsequently used in a parameter optimisation of the dynamic response of the ten introduced sandwich panels. The optimisation allows the identification of the most favourable multifunctional power structure. The experimental dynamic response of a manufactured multifunctional power panel is presented and compared to a conventional honeycomb panel for a successful evaluation of the introduced multifunctional approach. The results of this work show the ability of the presented multifunctional design to successfully combine the structural and power storage functions which makes the multifunctional power structure an excellent design approach for future space missions.

Microvibrations at frequencies between 1 and 1000 Hz generated by on-board equipment can propagate through a satellite's structure and hence significantly reduce the performance of sensitive payloads. This paper describes a Lagrange-Rayleigh-Ritz method for developing models suitable for the design of active control schemes. Here Loop Transfer Recovery based controller design methods are employed with this modeling strategy.

The AlSat-Nano mission is a joint endeavour by the UK and Algeria to build and operate a 3U CubeSat. The project was designed to provide training to Algerian students, making use of UK engineering and experience. The CubeSat was designed and built by the Surrey Space Centre (SSC) of the University of Surrey and hosts three UK payloads with operations run by the Algerian Space Agency (ASAL). The educational and CubeSat development were funded by the UK Space Agency (UKSA), whilst the UK payloads were self-funded. Launch and operations are funded by ASAL. This paper illustrates the development of the programme, the engineering of the satellite and the development of collaborative operations between the SSC and ASAL.

Inflatable technology is under continual development and advances in high strength fibres have pushed the limitations of these structures. This has led to their application in deploying large-aperture antennas, reflectors, solar sails and more recent large-scale inflatable spacecraft such as 'Transhab' and Genesis I and II. There is ongoing research into increasing the range of capabilities of these structures by enhancing the stiffness of the deployed structure. However, the lower stiffness and inherent flexibility of these inflatable structures also allows the structure to be morphed and controlled, which can be advantageous for a wide variety of applications. Initial research has previously been performed into integrating Shape Memory Alloys (SMA) with inflatable wings for Unmanned Aerial Vehicles (UAV) allowing airfoil optimization for various flight regimes. The aim of this work is to investigate the integration of shape memory alloys into inflatable structures for space applications. This article outlines ongoing research work at the University of Southampton into the field of morphing inflatable structures. Various commercially available shape memory alloys have been purchased and initially investigated to determine fundamental parameters such as maximum achievable force, power required, controllability and repeatability of the motion. These SMA's are then attached onto the side of a cantilevered inflatable cylindrical boom, using various attachment configurations, to investigate what tip deflections can be achieved for a voltage input. The tip deflection results of these configurations are presented along with an assessment of the technology and the areas of further research.

Flight and ground segment software in university missions is often developed only after hardware has matured sufficiently towards flight configuration and also as bespoke codebases to address key subsystems in power, communications, attitude, and payload control with little commonality. This bespoke software process is often hardware specific, highly sequential, and costly in staff/monitory resources and, ultimately, development time. Within Surrey Space Centre (SSC), there are a number of satellite missions under development with similar delivery timelines that have overlapping requirements for the common tasks and additional payload handling. To address the needs of multiple missions with limited staff resources in a given delivery schedule, computing commonality for both flight and ground segment software is exploited by implementing a common set of flight tasks (or modules) which can be automatically generated into ground segment databases to deliver advanced debugging support during system end-to-end test (SEET) and operations. This paper focuses on the development, implementation, and testing of SSC?s common software framework on the Stellenbosch ADCS stack and OBC emulators for numerous missions including Alsat-1N, RemoveDebris, SME-SAT, and InflateSail. The framework uses a combination of open-source embedded and enterprise tools such as the FreeRTOS operating system coupled with rapid development templates used to auto-generate C and Python scripts offline from ?message databases?. In the flight software, a ?core? packet router thread forwards messages between threads for inter process communication (IPC). On the ground, this is complemented with an auto-generated PostgreSQL database and web interface to test, log, and display results in the SSC satellite operations centre. Profiling is performed using FreeRTOS primitives to manage module behaviour, context, time and memory ? especially important during integration. This new framework has allowed for flight and ground software to be developed in parallel across SSC?s current and future missions more efficiently, with fewer propagated errors, and increased consistency between the flight software, ground station and project documentation.

In order to examine the dynamic response of spacecraft during launch, Coupled Loads Analyses (CLAs), which couple a Finite Element Model (FEM) of the spacecraft with a model of the launch vehicle, are performed to simulate critical flight events. For the CLA results to be trusted, it is necessary to first develop a high level of confidence in the spacecraft FEM. This confidence is achieved by conducting appropriate test-FEM correlation and update activities making use of data gathered during vibration testing of the physical hardware. One major point of concern is the containment of the correlation and update effort in terms of mode count/modal domain. As such, this work is concerned with the assessment of the effectiveness of various target mode selection criteria. Findings are presented for initial investigations conducted using FEM data for a large, unique, scientific spacecraft developed by the European Space Agency (ESA). The work presented herein is the initial stage, and a larger study would be required to draw conclusions on the most effective means of containing the modal domain for correlation and update activities to those natural frequencies/modes which are most likely to contribute significantly in response to flight event level loading conditions.

Test-analysis models are used in the validation of the nite element models of spacecraft structures. Here, a
probabilistic approach is used to assess the robustness of a system equivalent reduction expansion process based testanalysis
model when experimental and analytical modes contain different levels of inaccuracy. The approach is
applied to three spacecraft models, and Monte Carlo simulations were used to determine the sensitivity of the
normalized cross-orthogonality check to the system equivalent reduction expansion process reduced matrix. The
effect of parameters used in this reduction and the amount of inaccuracies that can be tolerated in the modes before
failing the normalized cross-orthogonality check were also determined. The results show that the probability to pass
the normalized cross-orthogonality check is highly determined by the number of modes used in the reduction. The
relation between capability of the nite element models to predict the frequency-response function and the quality of
the model validation determined using normalized cross-orthogonality check is also investigated, and it is observed
that the quantities are not always correlated. This study also shows that the sensor locations can be optimally chosen
using the system equivalent reduction expansion process reduced mass matrix, and this can increase the probability
to pass the normalized cross-orthogonality check.

A structure becomes a multifunctional power structure when in addition to meeting structural requirements it also performs functions associated with the electrical power system. With the structure performing these functions, some separate discreet components may no longer be required. Thus the parasitic structures that support them and the bus volume for these components are no longer required, reducing both mass and volume of the spacecraft. This paper focuses on the inclusion of commercial lithium polymer batteries into a sandwich panel which comprises the structure of a wing mounted solar array. It is shown that the thermal environment in earth orbit is hostile to the batteries. As such, a local thermal control system is required; with its authority targeted at preventing overcooling during eclipse. Phase change materials are proposed as a method to increase the thermal inertia of the structure by exploiting the latent heat. Through numerical simulation, it is shown that phase change materials are a relatively heavy solution. It is demonstrated that as the transition temperature rises, the amount of phase change material increases and that the optical properties of the structure can be altered to reduce the mass of phase change material required to more feasible levels.

Precision structures for space-based optical systems are typically subjected to brief periods of random vibration during the launch and ground testing phases. Such events pose a potential threat to the dimensional stability of such structures, which may be required to maintain positional tolerances on large optics in the low 10s of microns to meet optical performance requirements. Whilst there is an abundance of information in the literature on structural instability caused by hygrothermal cycling, there appears to have been little work done on the effects of random vibration. This issue has recently been addressed at RAL with a series of tests aimed at characterizing the behavior of dimensional instability in structures for high-resolution Earth-imaging cameras subject to random vibration. Firstly, a breadboard model of a typical "conventional" CFRP-based optical payload structure was produced and subjected to a range of environmental tests. The effects of random vibration were compared to those of other environmental stressors (such as thermal vacuum testing) and found to be significant. Next, controlled tests were performed on specific structural areas in order to assess the specific contributions of each area to overall instability. These tests made use of novel test setups and metrology techniques to assess the dimensional stability response of material samples and bolted joints to random vibration exposure. The tests were able to measure dimensional instability, characterize it over a series of tests of increasing vibration levels, and assess variability in results between identical samples. Finally, a predictive technique using a Finite Element Model with nonlinear kinematic hardening was produced. A time domain solution was obtained, using an analogy to Miner's Rule to determine load cycle amplitudes. This model correlated reasonably well with test results. This paper presents this program of work, and the results. It also proposes ways to minimize and mitigate dimensional instability due to random vibration by design, analysis and procedural means.

Microvibrations, at frequencies between 1 and 1000 Hz, generated by on board equipment, can propagate throughout a spacecraft structure and affect the performance of sensitive payloads. To investigate strategies to reduce these dynamic disturbances by means of active control systems, realistic yet simple structural models are necessary to represent the dynamics of the electromechanical system. In this paper a modeling technique which meets this requirement is presented, and the resulting mathematical model is used to develop some initial results on active control strategies. Attention is focused on a mass loaded panel subjected to point excitation sources, the objective being to minimize the displacement at an arbitrary output location. Piezoelectric patches acting as sensors and actuators are employed. The equations of motion are derived by using Lagrange's equation with vibration mode shapes as the Ritz functions. The number of sensors/actuators and their location is variable. The set of equations obtained is then transformed into state variables and some initial controller design studies are undertaken. These are based on standard linear systems optimal control theory where the resulting controller is implemented by a state observer. It is demonstrated that the proposed modeling technique is a feasible realistic basis for in-depth controller design/evaluation studies.

In this work, three techniques for the mathematical modelling of a piezoelectric actuated thin panel, namely the finite element method, a Lagrange Rayleigh-Ritz method, and a mechanical impedance-based method, are briefly presented. An accurate experimental implementation of a piezoelectric actuated simply-supported panel, whose dynamics have been simulated using the mathematical models, is described in detail. Since the differences between the results produced by the various mathematical models are very small, the accuracy of the experimental set-up is crucial. The results obtained via the numerical simulations are then compared with test results in order to assess the accuracy of the various modelling techniques.

For satellite applications the determination of the correct dynamic behaviour and in particular the structural damping is important to assess the vibration environment for the spacecraft subsystems and ultimately their capability to withstand the launch vibration environment. Therefore, the object of this investigation is to experimentally analyse a range of aluminium panel configurations to study the effect of joints on the damping of the complete structure. The paper begins with a full description of the experimental method used to accurately determine the modal loss factors for each of the panel configurations analysed. Nine different panels were used in the experimental tests, six of which incorporate lap joints variations. The joint parameters investigated include fastener type, bolt torque, fastener spacing, overlap distance and the effect of stiffeners. The damping results of ten different joint variants are presented for each of the first twelve modes of vibration. This data is directly compared to the damping factors of an equivalent monolithic panel. Various specific conclusions are made with respect to each of the joint parameters investigated. However, the primary conclusion is that the mode shape combined with the joint stiffness and joint location can be suggestive as to the likely magnitude increase of the modal loss factor.

Microvibrations of a RWA are usually studied in either hard-mounted or coupled conditions, although coupled wheel-structure disturbances are more representative than the hard-mounted disturbances. The coupled analysis method of the wheel-structure is not as well developed as the hard-mounted one. A coupled disturbance analysis method is proposed in this paper. One of the most important factors in coupled disturbance analysis - the accelerance or dynamic mass of the wheel is measured and results are validated with an equivalent FE model. The wheel hard-mounted disturbances are also measured from a vibration measurement platform particularly designed for this study. Wheel structural modes are solved from its analytical disturbance model and validated with the test results. The wheel-speed dependent accelerance analysis method is proposed.

In this paper, a full methodology to deal with microvibration predictions onboard satellites is described. Two important aspects are tackled: 1) the characterization of the sources with a pragmatic procedure that allows integrating into the algorithm the full effect of the sources, including their dynamic coupling with the satellite structure; 2) the modeling of the transfer function source receivers with a technique named in this paper as the Craig-Bampton stochastic method, which allows prediction of a nominal response and variations due to structural uncertainties as accurate as full Monte Carlo simulations but at a fraction of the computational effort. The paper then includes a statistical study of the data from the structural dynamic testing of the five identical craft of the Rapid-Eye constellation to set the magnitude of the uncertainties that should be applied in the analysis. Finally, the computational procedure is applied to the new high-resolution satellite SSTL-300-S1 and the predictions compared with the experimental results retrieved during the physical microvibration testing of the satellite, which was carried out at the Surrey Satellite Technology Limited facilities in the United Kingdom.

Since the beginning of the space era, a significant amount of debris has progressively been generated in space. Active Debris Removal
(ADR) missions have been suggested as a way of limiting and controlling future growth in orbital space debris by actively deploying
vehicles to remove debris. The European Commission FP7-sponsored RemoveDebris mission, which started in 2013, draws on the
expertise of some of Europe?s most prominent space institutions in order to demonstrate key ADR technologies in a cost effective
ambitious manner: net capture, harpoon capture, vision-based navigation, dragsail de-orbiting.
This paper provides an overview of some of the final payload test results before launch. A comprehensive test campaign is
underway on both payloads and platform. The tests aim to demonstrate both functional success of the experiments and that the
experiments can survive the space environment. Space environmental tests (EVT) include vibration, thermal, vacuum or thermalvacuum
(TVAC) and in some cases EMC and shock. The test flow differs for each payload and depends on the heritage of the
constituent payload parts. The paper will also provide an update to the launch, expected in 2017 from the International Space Station
(ISS), and test philosophy that has been influenced from the launch and prerequisite NASA safety review for the mission.
The RemoveDebris mission aims to be one of the world?s first in-orbit demonstrations of key technologies for active debris
removal and is a vital prerequisite to achieving the ultimate goal of a cleaner Earth orbital environment.

During launch, a spacecraft undergoes loads ranging from quasi-static to highly transient or harmonic low frequency events, from higher frequency shock loads to acoustic excitations. In order to reproduce such a dynamic diversity, fixed base sinusoidal tests, wide band acoustic loading and different regimes of shock testing are implemented in the test facilities. In this article, the main focus is on fixed base sinusoidal tests, fundamental for a number of reasons, including demonstrating that the satellite can withstand the low frequency dynamic environment and validating the mathematical model which will then be also used for coupled load analysis purposes. For the latter, a post-test correlation process is carried out and the basic assumption is trusting the experimental results obtained from shaker testing. In reality, some of these assumptions (e.g. ?infinitely? stiff boundary and inertial properties of the shaker) are not correct, as for the kind of applications treated in this article experimental results are significantly affected by boundary flexibilities, modes of the shaker/head expander and non-perfect implementation of the control algorithm in the electronic hardware. In the last decade, there has been a growing interest in virtual testing, with the long-term view to use simulation as substitute for the majority of testing, but currently under investigation for pre-test response predictions and post-test correlation. Here, the satellite is mathematically modelled along with the shaker and the control system. In this article, in particular, a simulation capability of longitudinal closed loop control simulation of the ESA electrodynamic shaker (QUAD) flexible body coupled with a test specimen (Bepi Colombo) flexible model is developed. This shows how significant the differences are when looking at the analytical results from two different perspectives (standard Finite Element Analysis and Virtual Testing implementation). The focus of this article is specifically on post-test correlation: correlation methods are used for both procedures and results show significant improvements when the satellite Finite Element Model undergoes the virtual testing approach.

This paper describes the scalability analysis of bistable Carbon Fibre Reinforced Plastic (CFRP) tubes for space applications, with the aim of attaining a better understanding of the scaling laws of Bistable Reeled Composite (BRC) tubes. BRCs with substantially higher natural frequency are designed. The application for this work is a deployable solar array, which uses two BRC tubes to support a membrane containing flexible photovoltaic cells. Novel types of bistable tubes with stepped thickness changes, tapered diameter and reduced included angle are proposed to improve the natural frequency. Finite Element (FE) modelling and experimental verification have been used to study the vibration characteristics of the proposed BRC tubes. An FE model is combined with an optimization loop to improve the natural frequency with respect to the fibre angles within the laminate of the bistable tubes. The results demonstrate that the introduction of step changes in laminate thickness at certain locations, and careful selection of fibre angles can significantly improve the natural frequency.

This paper addresses the characterisation and analysis of a 2-collinear-DoF strut with embedded electromagnetic shunt dampers. The use of a negative resistance in the shunt circuit has been proved to considerably enhance the damping performance of this kind of electromagnetic dampers. The analytical model is reported and the theoretical results are compared with other damping methods. This work demonstrates the feasibility of achieving a remarkable decay rate of -80 dB/decade with a device that is smaller than previously-presented active struts and does not require complex electronics to operate.

Micro-vibration on board a spacecraft is an important issue that affects payloads requiring high pointing accuracy. Although isolators have been extensively studied and implemented to tackle this issue, their application is far from being ideal due to the several drawbacks that they present, such as limited low-frequency attenuation for passive systems or high power consumption and reliability issues for active systems. In the present study, a novel 2-collinear-DoF strut with embedded electromagnetic shunt dampers (EMSD) is modelled, analysed and the concept is physically tested. The combination of high-inductance components and negative-resistance circuits is used in the two shunt circuits to improve the EMSD micro-vibration mitigation and to achieve an overall strut damping performance that is characterised by the elimination of the resonance peaks and a remarkable FRF final decay rate of ?80 dB dec?1. The EMSD operates without requiring any control algorithm and can be comfortably integrated on a satellite due to the low power required, the simplified electronics and the small mass. This work demonstrates, both analytically and experimentally, that the proposed strut is capable of producing better isolation performance than other well-established damping solutions over the whole temperature range of interest.

Driven by the increasingly stringent stability requirement of some modern payloads (e.g. the new generations of optical instruments) the issue of accurate spacecraft micro-vibration modeling has grown increasingly important. In this context micro-vibrations are low level mechanical disturbances occurring at frequencies from a few Hertz up to 1000 Hz. As the frequency content of these phenomena extends beyond the first few modal frequencies, FEA predictions become less accurate and alternative methods have to be considered. Other modeling and analysis techniques have been investigated and applied to vibration problems (Stochastic Finite Element Method (e.g. Monte Carlo Simulation), Statistical Energy Analysis (well-established method for high frequency ranges) and the Hybrid FE-SEA), with the aim of investigating medium and high frequency behavior. This work is part of a project whose aim is to establish appropriate procedures for the modeling and analysis of micro-vibration and validate these procedures against experimental data. All the methods cited above are implemented in this study and compared with experimental results, in order to assess the performance of the various methodologies for micro-vibration problems, covering the whole frequency range up to 1000 Hz. Some comparisons between experimental and computational results are performed using the MAC. Some other analyses, like linearity, reciprocity or effect of the harness are also described. The bench work model that has provided the experimental data is the satellite platform SSTL 300 and this paper outlines these related test campaigns.

In recent years, driven by the increasingly stringent stability requirements imposed by some satellites? payloads (e.g., the new generation of optical instruments), the issue of accurate onboard spacecraft microvibration modeling has attracted significant interest from engineers and scientists. This paper investigates the microvibration-induced phenomenon on a cantilever-configured reaction wheel assembly including sub- and higher harmonic amplifications due to modal resonances and broadband noise. A mathematical model of the reaction wheel assembly is developed and validated against experimental test results. The model is capable of representing each configuration in which the reaction wheel assembly will operate, whether it is hard mounted on a dynamometric platform or suspended free?free. The outcomes of this analysis are used to establish a novel methodology to retrieve the dynamic mass of the reaction wheel assembly in its operative range of speeds. An alternative measurement procedure has been developed for this purpose, showing to produce good estimates over a wide range of frequencies using a less complex test campaign compared with typical dynamic mass setups. Furthermore, the gyroscopic effect influence in the reaction wheel assembly response is thoroughly examined both analytically and experimentally. Finally, to what extent the noise affects the convergence of the novel approach is investigated.

Thin metal-polymer laminates make excellent materials for
use in inflatable space structures. By inflating a stowed envelope
using pressurized gas, and by increasing the internal
pressure slightly beyond the yield point of the metal films,
the shell rigidizes in the deployed shape. Structures constructed
with such materials retain the deployed geometry
once the inflation gas has either leaked away, or it has been
intentionally vented. For flight, these structures must be initially
folded and stowed. This paper presents a numerical
method for predicting the force required to achieve a given
fold radius in a three-ply metal-polymer-metal laminate and
to obtain the resultant springback. A coupon of the laminate
is modeled as a cantilever subject to an increasing tip
force. Fully elastic, elastic-plastic, relaxation and springback
stages are included in the model. The results show good
agreement when compared with experimental data at large
curvatures.

The severe mechanical environment that electronic components experience during spacecraft launch necessitates that their failure probability be assessed. One possible approach is to create an accurate model of the Printed Circuit Boards (PCBs) dynamic response; subsequently the failure probability can be determined by comparing the response model with corresponding failure criteria for the electronic components. In principle the response model can be achieved by a very detailed Finite Element (FE) model of the PCB which would include the mass and stiffness of all components present on the PCB. Unfortunately this approach requires an excessive effort; therefore it is rarely pursued by the designer. Past research has shown that assumptions can be made about the mass and stiffness that allow simpler models to be created that still achieve appropriate levels of accuracy. However, the accuracy of these simplified models has not yet been quantified over a range of possible design cases. This paper will quantify how increasing levels of modelling simplifications decrease the accuracy of PCB FE models.

Coupled Loads Analyses (CLAs), using finite element models (FEMs) of the spacecraft and launch vehicle to simulate critical flight events, are performed in order to determine the dynamic loadings that will be experienced by spacecraft during launch. A validation process is carried out on the spacecraft FEM beforehand to ensure that the dynamics of the analytical model sufficiently represent the behavior of the physical hardware. One aspect of concern is the containment of the FEM correlation and update effort to focus on the vibration modes which are most likely to be excited under test and CLA conditions. This study therefore provides new insight into the prioritization of spacecraft FEM modes for correlation to base-shake vibration test data. The work involved example application to large, unique, scientific spacecraft, with modern FEMs comprising over a million degrees of freedom. This comprehensive investigation explores: the modes inherently important to the spacecraft structures, irrespective of excitation; the particular ?critical modes? which produce peak responses to CLA level excitation; an assessment of several traditional target mode selection methods in terms of ability to predict these ?critical modes?; and an indication of the level of correlation these FEM modes achieve compared to corresponding test data. Findings indicate that, although the traditional methods of target mode selection have merit and are able to identify many of the modes of significance to the spacecraft, there are ?critical modes? which may be missed by conventional application of these methods. The use of different thresholds to select potential target modes from these parameters would enable identification of many of these missed modes. Ultimately, some consideration of the expected excitations is required to predict all modes likely to contribute to the response of the spacecraft in operation.

Bistable composite shells patented as Bistable Reeled Composite (BRC) booms have the potential to be used as lightweight structural elements for a number of space applications. This paper details an approach to increase the natural frequency and stiffness of BRCs. The motivation for this research is the desire to increase the scalability of a flexible "roll-up" solar array which, in its deployed state, consists of two cantilevered BRCs supporting a flexible Photo Voltaic (PV) cell covered blanket between them. A Finite Element (FE) numerical model is combined with a nonlinear constrained optimization to maximize the natural frequency of BRC booms with respect to the fiber orientation angles and ply discontinuity locations. The results demonstrate that careful selection of the fiber orientation angles and the location of step thickness variations can significantly optimize the natural frequency. Experimental verification of the vibration characteristics of optimized BRC booms has also been conducted. Finally, stability analysis of the optimized BRC booms under bending has been carried out using FE simulation to quantify the Maximum Rotational Acceleration (MRA) that they can take before failure.

Mathematical finite element models (FEMs) of spacecraft are relied upon for the prediction of loads experienced during launch and flight events. It is essential that the spacecraft is able to survive the launch environment without sustaining damage which could inhibit its ability to carry out its mission. Therefore, ensuring that these FEMs give a realistic representation of the physical spacecraft structural dynamics is an important task. To achieve a high level of confidence in the FEM in question, a correlation activity is conducted. This is the process of applying various metrics to compare computational results, from analysis of the FEM, with corresponding data derived from measurements taken of the physical hardware during vibration testing. Subsequently, updates are applied to the FEM where necessary to achieve an acceptable level of correlation.
It is possible for spacecraft FEM correlation exercises to take a considerable amount of time and effort without necessarily achieving an appreciable improvement in the final FEM. As such, this project has been conducted to address the need to ensure that the procedures being applied are as effective and efficient as possible. Various aspects of the spacecraft FEM correlation process have been investigated separately, and interactions between the different stages in the process have also been considered. Two large, unique, scientific spacecraft have been used as example applications in order to carry out these studies. As well as making use of computational results from the spacecraft FEMs, this project has also included comparisons to the results from the corresponding base-shake sine-sweep test campaigns conducted on these structures.
A number of noteworthy, and industrially beneficial, findings relating to the effectiveness of the spacecraft FEM correlation process have resulted from these studies: the most appropriate techniques of modal parameter estimation for the considered spacecraft applications have been established; the potential benefits and relative merits of different pre-test sensor placement procedures have been explored; inaccuracies introduced through the use of a commonly applied FEM reduction method have been demonstrated and a superior alternative identified. In addition, the efficiency of the correlation and update process has also been addressed. This has mainly been achieved through investigations concerning the applicability of commonly used target mode selection criteria to spacecraft applications, and the potential benefits of a less widely applied method which takes into consideration the expected loading scenarios to be experienced by the considered structures.

Deployable booms are an essential part of the deployable structures family used in space. They can be stowed in a coiled form and extended into a rod like structure in an action similar to that of a carpenter?s tape measure. ?Blossoming? is a failure mode that some boom deployers experience where the booms uncoil within the deployer instead of extending. This paper develops a method to predict the force that a boom can exert before blossoming occurs by using the strain energy stored in the coiled boom and in the compression springs. An experimental apparatus is used to gain practical results to compare to the theory.

Nowadays, a technology demonstrator platform popular amongst the research community given their relatively low cost and short development time are cubesats. Nevertheless, cubesats are by definition nano-satellites of small volume and mass, and therefore, they traditionally only allowed very limited sizes of any expandable structure onboard with final deployed areas in the order of a few square meters. This conflicts with the large areas required for efficient solar sails, making the demonstration of this exotic concept bound to more expensive missions with a dedicated launch. The applications that will be discussed throughout the thesis will be: three-axis stabilised solar sailing with a "rigid" support structure; and drag assisted deorbiting of a large host craft using a solar sail. Both of these applications still need validation in space, especially for Earth-bound missions.

The main goal of this research effort is thus to satisfy the need of available deployable booms for their use on systems of unprecedented mass per unit area with cubesat-like mission constraints that will ultimately place more trust in gossamer concepts. For this, two novel rollable booms and their deployment mechanisms have been developed, one based on metallic tape-springs and the other on bistable composite slit tubes. Analyses and tests confirmed that the former boom has scalability problems related to stowage-induced boom axial curvature, and coil blossoming management. Reliable sail deployments of a 4 x 4 m^2 sail were achieved with them. The latter boom design solves previous scalability problems of bistable composite booms. The ground demonstrator tested deploys reliably a 5 x 5 m^2 sail, with the current compact boom design shown to be efficiently scalable for 100 m^2 class sails. To enable even larger sails with the bistable booms developed, a novel architecture named the completely stripped solar sail has been proposed. A simple experiment demonstrated the beneficial effect that dividing the sail into sets of parallel strips and using a continuous sail-boom attachment suspension configuration has towards scalability of the concept.

A new structural characterisation programme developed means by which to characterise the slender booms properties. In addition, the test results validated and/or updated the imperfection seeded finite element models produced. These models are ultimately utilised in high-fidelity predictions of the performance of the solar sail booms under the established operational loads, as well as in the scalability analyses of the sail concepts proposed with them.

Lastly, the first gossamer sail-based deorbiting system in it class, developed for medium mass (

This paper describes progress towards developing design guidelines for a number of composite bonded joints in aerospace applications. The premise of a universal failure criterion is impractical given the number of adherend-adhesive configurations and layups available. However, for a finite number of joint configurations, design rules can be developed based on experimental test data and detailed finite element (FE) modelling. By using these techniques rather than the traditional overly conservative knock down factors, more of the performance of composite bonded joints can be accessed. The work presented here experimentally studied the effect of the substrate layup, adhesive type and adhesive thickness on double-lap joint (DLJ) strength. The corresponding failure surfaces were analysed and failure modes identified. Following this, detailed FE models were developed to identify the trends associated with altering joint parameters. Finally, the stresses and strains within the adhesive and substrate were analysed at the joints respective failure loads to identify critical parameters. These parameters can provide an insight as to the stress state of the joint at failure or near failure loads, and hence its true performance.

The stringent stability requirements imposed by advanced, high-resolution payloads have
produced an increased interest in the development of better-performing micro-vibration
isolators. Several devices aimed at mitigating micro-vibrations have been studied and
implemented, but their application is still far from being ideal due to the several drawbacks
that they present, such as limited low-frequency attenuation for passive systems or high
power consumption and reliability issues for active systems.
This research focuses on the modelling and testing of Electromagnetic Shunt Dampers
(EMSD) characterised by the use of negative impedance converter circuits. An electromagnetic
damper is a self-excited device that exploits the interaction between a moving
magnetic field and a conductive material to provide a reaction force to the applied motion.
An EMSD presents several advantages, but the high ratio of system mass over damping force
produced has limited its application in space missions. The use of a negative resistance can
considerably lower this ratio since it produces an overall reduction of the circuit resistance
that results in an increase of the induced current in the closed circuit and thus the damping
performance.
In this thesis, the development of a multiphysics, multi-parametric model of an EMSD is
presented and accurately corroborated by an extensive test campaign. This damper can be
classified as a semi-active damper since the negative resistance circuit does not require any
control algorithm to operate. In terms of damping performance, this research demonstrates
that an EMSD applied to a 1-DoF system is capable of behaving, throughout the whole
temperature range of interest, like a 2nd-order mechanical filter in which the resonance
peak is eliminated and the roll-off slope is -40 dB/dec. Additionally, the proposed EMSD is
characterised by low required power, simplified electronics and small device mass that could
allow it to be comfortably integrated on a satellite.
This study presents also a possible novel 2-collinear-DoF system design with embedded
EMSDs. This isolator is capable of achieving a remarkable final decay rate of -80 dB/dec
while completely eliminating the two resonance peaks due to the high attenuation performance
of the dampers. Moreover, other aspects of the proposed 2-collinear-DoF system
are investigated in order to assess not only the damping performance but also its features at
system level. This work demonstrates that the fundamental advantages of this system can
make it a viable, competitive alternative to other actively controlled struts.

Deployable structures play an important role in space applications as they minimise the volume required by large structures such as antennas, solar panels, reflectors or de-orbiters. A low cost and mass option, relies on the use of airtight inflatable structures. Over the years several rigidization methods have been developed, each with their strengths and weaknesses, however due to the simplicity of aluminium based metal-polymer laminates, this class of shells have been successfully flown on a series of legacy missions. Metal polymer laminates are typically three-ply constructions where two foils of ductile annealed aluminium sandwich a polymer core. Structures such as sphere and columns may be constructed from flat sheets of material. The envelopes are then packaged. Pressurised gas, typically nitrogen is realised into the envelope to achieve deployment. To rigidize the structure the pressure is further increased to a value slightly higher than the yield point of the metal foils. By conducting repeated rigidization experiments it was observed that residual fold creases remain present in metallic shell. As metal laminates rely on the structural integrity of the shell for strength, it is important that the extent of the initial imperfections is known. The collapse load of laminate columns is significantly reduced by this effect. If care is taken during packaging and construction of these structures, packaging residual creases remain the largest source of imperfections.

To observe closely the folding process, a 3D laser and SEM images have been taken at various steps during folding. To understand this mechanism these results were compared against the results from literature. It has been found that for a `Z' folded column the longitudinal creases flatten more than the circumferential creases.

A numerical model has been derived for the elastic-plastic bending and springback of a metal film and metal-polymer-metal laminate. In the presented work this approach replicates the introduction of a typical `V' fold and relaxation once the load has been removed. The system of differential equations was solved in MATLAB using ode45. To simplify the analysis a bilinear stress-strain profile with plane strain has been attributed to the metal film. The results have been validated with good agreement against experimental results and FEA analysis conducted in ABAQUS. Two three ply aluminium-polymer-aluminium flight ready laminates have been used as the experimental benchmark. The derived model may be adapted for different laminate configurations. It is known that it is difficult to quantify the mechanical properties of thin aluminium films, in particular the Young?s modulus. Several results from literature are discussed and the solutions proposed is outlined.

The lessons learn from this research project have been applied to the development of a novel rigidizable aluminium-BoPET based deployable structure. The structure consists of six laminate booms to connect to form a cuboid structure with a cross-sectional are of 0.5 m2. The structure and support systems were design to occupy the volume of single CubeSat Unit. The deployable will be flown on the RemoveDebris ADR technology demonstrator.

Coiled deployable booms have been used extensively in space and are a large part of the deployable
space structures family. They have a wide variety of uses such as the deployment of instruments,
gravity-gradient stabilisation masses and more recently solar sails. Most deployable booms are
similar to a carpenter's tape measure in the way they are coiled in a retracted condition and then
deploy to form the boom structure. There have been many developments in the optimisation of
boom properties in the deployed state, by using different shape cross sections and by using different
materials. The ýfirst metal tape spring booms have developed into the more modern booms with
a variety of cross sections. One aspect that is common to all booms is the coiling and uncoiling
process and the difficulties associated with this. Blossoming, where the boom starts to uncoil within
the boom deployer, can lead to the jamming of the mechanism. The reasons behind blossoming
have not been thoroughly investigated, leaving designers of booms, and boom housing mechanisms
to try and mitigate this problem themselves, often by trial and error. This work investigates boom
blossoming with the aim of better understanding the underlying mechanics so that more effective
deployment systems can be designed in the future.
A method is developed that uses the strain energy stored in coiled booms to ýfind the maximum
tip force that can be achieved before blossoming occurs. This method is also used to investigate
the central spindle torque during blossoming. The effects that the coil geometry and the friction
between the layers of the coiled booms have on blossoming are also investigated.
The theory developed should enable the designers of tape spring deployers to estimate the tip force
and central spindle torque of a tape spring boom in the design phase of projects and reduce the
reliance on trial and improvement type testing once deployers have already been built.

Spacecraft overtesting is a long running problem, and the main focus of most attempts to reduce it has been to adjust the base vibration input (i.e. notching). Instead this paper examines testing alternatives for secondary structures (equipment) coupled to the main structure (satellite) when they are tested separately. Even if the vibration source is applied along one of the orthogonal axes at the base of the coupled system (satellite plus equipment), the dynamics of the system and potentially the interface configuration mean the vibration at the interface may not occur all along one axis much less the corresponding orthogonal axis of the base excitation.
This paper proposes an alternative testing methodology in which the testing of a piece of equipment occurs at an offset angle. This Angle Optimisation method may have multiple tests but each with an altered input direction allowing for the best match between all specified equipment system responses with coupled system tests. An optimisation process that compares the calculated equipment RMS values for a range of inputs with the maximum coupled system RMS values, and is used to find the optimal testing configuration for the given parameters.
A case study was performed to find the best testing angles to match the acceleration responses of the centre of mass and sum of interface forces for all three axes, as well as the von Mises stress for an element by a fastening point. The angle optimisation method resulted in RMS values and PSD responses that were much closer to the coupled system when compared with traditional testing. The optimum testing configuration resulted in an overall average error significantly smaller than the traditional method. Crucially, this case study shows that the optimum test campaign could be a single equipment level test opposed to the traditional three orthogonal direction tests.

Efficient vibroacoustic response prediction on complex structures, such as spacecraft, represents
a challenging task, even for the computers and numerical techniques of today. This is particularly
evident in the mid-frequency range, where structures begin exhibiting chaotic behaviour, rendering
element-based techniques inefficient or unreliable.
In this article, an efficient random formulation for reduced finite element method (FEM) models is
proposed, such that Monte Carlo simulations can be carried out robustly within practically acceptable
timeframes. The introduced novel non-parametric stochastic FEM is inherently compatible
with various existing component mode synthesis techniques. It is particularly well adapted to use
with popular modal reduction approaches, such as the Craig-Bampton method. The mathematical
framework for the method is outlined, enabling the deterministic reduced matrices to be robustly
perturbed at the subsystem level. Properties, such as matrix positive-(semi)definiteness, mean
system eigenvalues, and representation accuracy are preserved. This new stochastic FEM is validated
against a full parametric Monte-Carlo simulation and test data of a real spacecraft structure,
establishing its reliability and computational efficiency.
In the proposed coupled FEM-BEM approach, the acoustic domain is modelled with hierarchical
matrix accelerated collocation BEM. This alleviates the memory requirements for the large, dense
BEM matrices, and the need for spatial discretisation of acoustic FEM. The full implementation
is outlined for a simple geometry discretised with high a density mesh, showing consistent convergence
of the employed iterative solver.

In this paper, the mathematical framework for a computationally
efficient stochastic finite element method (FEM)
is outlined. It is devised for a range of applications in
structural dynamics, where uncertainties need to be reliably
dealt with in the context of reduced model formulations.
It allows random mass and stiffness matrices
to be robustly generated at the subsystem level in component
mode synthesis (CMS) applications. The technique
is validated for the particularly challenging case
of mid-frequency FEM-FEM vibroacoustic analysis of a
spacecraft structure. Results are compared against both
test data and full parametric Monte-Carlo simulation. Finally,
the method?s applicability to coupled vibroacoustic
problems utilising hierarchical matrix boundary element
method (BEM) acoustic formulations is evaluated.

Disturbances generated by reaction wheels on board the
spacecraft are among the most
substantial
. Hence they
play a crucial role when microvibration budget
has to be
assessed. This paper aims at characterising the effects of
RW on the structure by focusing on the format of the
disturbance input matrix of these components. In
particular the case of single and multiple wheel
accounted for. In the first
one
the
responses are evaluated
at some specific locations of the reaction wheel where
their disturbance is amplified, i.e. harmonics. In the
second case a more realistic scenario is considered with
several wheels to be characterised and the effects of
neglecting
some terms of the disturbance input matrix are
discussed. Finally a sensitivity analysis is carried out to
quantify in which extent changes in the input matrix can
alter the response. A preliminary methodology is then
suggested to characterise a large num
ber of wheels.

Due to constantly increasing requirements for more precise and high-resolution instrumentations,
microvibration prediction represents an issue of growing importance. Hence the need
of reliable analysis tools which can evaluate microvibrations effects efficiently. This paper
describes how to tackle the issue of structural uncertainties in microvibration predictions. In
particular, uncertainties related to the microvibration sources are analysed as well as those
linked to the modelling of the structure. A methodology to define the worst case of vibration
produced by on board sources is presented and compared to experimental data. Additionally,
an approach to quantify the uncertainties in the Finite Element model is also described.

This thesis describes progression towards developing an enhanced design methodology for laminated composite bonded joints in aerospace applications. The premise of a universal failure criterion is impractical given the number of adhesive-adherend configurations available. However, for a finite number of joint configurations, design rules can be developed based on experimental test data and detailed finite element modelling. By using these techniques rather than the traditional, overly conservative knock-down factors, more of the performance of composite bonded joints can be accessed. While complex damage modelling techniques are available, the additional material data and analysis time required renders them not suitable for the vast majority of time-sensitive industrial applications.
Initially, the work presented in this thesis experimentally studied the effect of the substrate material, substrate layup, adhesive material and adhesive thickness on several laminated composite bonded joint configurations. The corresponding failure surfaces were extensively analysed and failure modes identified. Following this, detailed FE models were developed to identify the trends associated with altering joint parameters. Finally, the stresses and strains within the adhesive and substrate were analysed at each joint?s respective failure loads to identify critical parameters, which would later be used to develop a Critical Parameter Method for evaluating joint performance.
Once these parameters were consolidated, they were validated against a unique set of joints. The critical parameter approach was able to predict joint strength with an average error of 26% compared experimental strength. Traditional FE criterions presented an average error of 61% compared to experimental strength. After further consolidation, joint strength prediction reduced to within 3% of experimental strength using the Critical Parameter Method, representing a substantial improvement in predictive capabilities.

In recent years extremely small satellites have been developed in response to trends in the space industry to achieve more for less cost. Extremely lightweight and efficiently packaged deployable structures are essential for achieving large-scale applications including communication antennas, solar arrays, and in recent years, deorbiting drag-sails.

This thesis is motivated for developing novel deployable helical antennas for space-based maritime surveillance. The helical antenna technology provides packaging efficiency and radio frequency characteristics superior to the latest efforts of international research groups. To achieve this, the research presented focuses on developing the proven bistable composite slit tube (BCST) deployable technology. These are open-section tubular structures which can be deployed and rolled up into a compact coil, analogous to a tape measure, but do not require constraint to remain stowed. This behaviour is referred to as bistability and enables lightweight and relatively simple deployable structures for spacecraft applications.

New forms of BCST are modelled through the introduction of additional curvatures, manufactured and described in this work with two new subcategorisations established: toroidal and helical. Toroidal BCSTs are doubly curved with both principal curvatures initially non-zero in the deployed stress-free state. Helical BCSTs are doubly curved and twisted out-of-plane. Investigations into the effects of geometrical parameters and laminated composite material properties on the bistable coils of both types are presented. The results provide an understanding of bistable behaviour in new forms of BCST previously neglected in the literature, which is almost exclusively focused on straight forms. As a consequence of this research, new deployable structure technologies are envisaged in the areas of compact terrestrial shelters and small satellite communications.

The InflateSail CubeSat, designed and built at the Surrey Space Centre (SSC) at the University of Surrey, UK, for the Von Karman Institute (VKI), Belgium, is one of the technology demonstrators for the QB50 programme. The 3.2 kilogram InflateSail is ?3U? in size and is equipped with a 1 metre long inflatable boom and a 10 square metre deployable drag sail.
InflateSail's primary goal is to demonstrate the effectiveness of using a drag sail in Low Earth Orbit (LEO) to dramatically increase the rate at which satellites lose altitude and re-enter the Earth's atmosphere. InflateSail was launched on Friday 23rd June 2017 into a 505km Sun-synchronous orbit. Shortly after the satellite was inserted into its orbit, the satellite booted up and automatically started its successful deployment sequence and quickly started its decent. The spacecraft exhibited varying dynamic modes, capturing in-situ attitude data throughout the mission lifetime. The InflateSail spacecraft re-entered 72 days after launch.
This paper describes the spacecraft and payload, and analyses the effect of payload deployment on its orbital trajectory. The boom/drag-sail technology developed by SSC will next be used on the RemoveDebris mission, which will demonstrate the applicability of the system to microsat deorbiting.

The InflateSail (QB50-UK06) CubeSat, designed and built at the Surrey Space Centre (SSC) for the Von Karman
Institute (VKI), Belgium, was one of the technology demonstrators for the European Commission?s QB50
programme. The 3.2 kg 3U CubeSat was equipped with a 1 metre long inflatable mast and a 10m2
deployable drag
sail. InflateSail's primary mission was to demonstrate the effectiveness of using a drag sail in Low Earth Orbit (LEO)
to dramatically increase the rate at which satellites lose altitude and re-enter the Earth's atmosphere and it was one of
31 satellites that were launched simultaneously on the PSLV (polar satellite launch vehicle) C-38 from Sriharikota,
India on 23rd June 2017 into a 505km, 97.44o
Sun-synchronous orbit.
Shortly after safe deployment in orbit, InflateSail automatically activated its payload. Firstly, it inflated its metrelong
metal-polymer laminate tubular mast, and then activated a stepper motor to extend four lightweight bi-stable
rigid composite (BRC) booms from the end of the mast, so as to draw out the 3.1m x 3.1m square, 12mm thick
polyethylene naphthalate (PEN) drag-sail. As intended, the satellite immediately began to lose altitude, causing it to
re-enter the atmosphere just 72 days later ? thus successfully demonstrating for the first time the de-orbiting of a
spacecraft using European inflatable and drag-sail technologies.
The InflateSail project was funded by two European Commission Framework Program Seven (FP7) projects:
DEPLOYTECH and QB50. DEPLOYTECH had eight European partners including DLR, Airbus France, RolaTube,
Cambridge University, and was assisted by NASA Marshall Space Flight Center. DEPLOYTECH?s objectives were
to advance the technological capabilities of three different space deployable technologies by qualifying their
concepts for space use. QB50 was a programme, led by VKI, for launching a network of 50 CubeSats built mainly by
university teams all over the world to perform first-class science in the largely unexplored lower thermosphere.
The boom/drag-sail technology developed by SSC will next be used on a third FP7 Project: RemoveDebris,
launched in 2018, which will demonstrate the capturing and de-orbiting of artificial space debris targets using a net
and harpoon system. This paper describes the results of the InflateSail mission, including the observed effects of
atmospheric density and solar activity on its trajectory and body dynamics. It also describes the application of the
technology to RemoveDebris and its potential as a commercial de-orbiting add-on package for future space missions.

The size of any single spacecraft is ultimately limited by the volume and mass constraints of currently available
launchers, even if elaborate deployment techniques are employed. Costs of a single large spacecraft may also be
unfeasible for some applications such as space telescopes, due to the increasing cost and complexity of very large
monolithic components such as polished mirrors.

The capability to assemble in-orbit will be required to address missions with large infrastructures or large
instruments/apertures for the purposes of increased resolution or sensitivity. This can be achieved by launching
multiple smaller spacecraft elements with innovative technologies to assemble (or self-assemble) once in space and
build a larger much fractionated spacecraft than the individual modules launched.

Up until now, in-orbit assembly has been restricted to the domain of very large and expensive missions such as space
stations. However, we are now entering into a new and exciting era of space exploitation, where new mission
applications/markets are on the horizon which will require the ability to assemble large spacecraft in orbit. These
missions will need to be commercially viable and use both innovative technologies and small/micro satellite
approaches, in order to be commercially successful, whilst still being safety compliant. This will enable
organisations such as SSTL, to compete in an area previously exclusive to large commercial players. However, inorbit
assembly brings its own challenges in terms of guidance, navigation and control, robotics, sensors, docking
mechanisms, system control, data handling, optical alignment and stability, lighting, as well as many other elements
including non-technical issues such as regulatory and safety constraints. Nevertheless, small satellites can also be
used to demonstrate and de-risk these technologies.

In line with these future mission trends and challenges, and to prepare for future commercial mission demands, SSTL
has recently been making strides towards developing its overall capability in ?in-orbit assembly in space? using
small satellites and low-cost commercial approaches. This includes studies and collaborations with Surrey Space
Centre (SSC) to investigate the three main potential approaches for in-orbit assembly, i.e. deployable structures,
robotic assembly and modular rendezvous and docking. Furthermore, SSTL is currently developing an innovative
small ~20kg nanosatellite (the ?Target?) as part of the ELSA-d mission which will include various rendezvous and
docking demonstrations. This paper provides an overview and latest results/status of all these exciting recent in-orbit
assembly related activities.

This article examines the new practice of Virtual Shaker Testing (VST), starting from its motivation to its practical implementations and future possible implications. The issues currently experienced during large satellites? vibration testing are discussed, examining practical examples that highlight the coupling existing between the item under test and facility, and that are the basis for the motivation behind the new methodology (i.e. VST). VST is proposed as a way to bypass some of these issues, and here its use as a pre and post shaker test tool is discussed. In the article VST is applied to real test cases (Airbus? large spacecraft Bepi Colombo, built for the European Space Agency's first mission to Mercury), showing computations and real physical test data to illustrate the advantages of the methodology. These are mostly in terms of de-risking of the physical test campaigns (due to the capability to simulate realistically the future physical test thus reducing the probability of aborts and stops during the runs), and an improvement of the quality of the correlation process and related FEM update (resulting from the capability to separate the dynamics of the satellite from the effects of the test equipment); ultimately providing a tool to address questions arising from test response observations, which are many. This tool also offers the possibility to improve vibration testing using 6 DOF facilities. The article is concluded articulating a possible way forward to take maximum advantage of the new methodology, drawing a parallel with the current Satellite/Launch Vehicle Coupled Load Analysis cycles, and proposing a different design and validation philosophy.

It is well documented that reaction wheels are among the most significant microvibration sources in space applications. These components, despite being nominally identical, can show differences in the generated signals due to manufacturing imperfections in their internal elements, such as ball bearing, internal and external race. In this article a methodology to account for those variations in microvibration predictions is proposed, aiming at generating a disturbance input matrix that encompasses the effects of a family of reaction wheels. With such a tool, it is possible to provide a more accurate microvibration budget at an early stage of the mission, reducing the uncertainty margin usually applied to quantify reaction wheel effects on the structure. As a consequence better designs are produced faster and cheaper. This allows for more flexibility in the mission design and reduces the degree of uncertainties in the predictions. Furthermore, it is shown that the proposed approach is able to characterise the effects of the entire family of wheels by considering only a limited number. The methodology is validated by assessing the microvibration excitation on different structures, including a real space structure with various reaction wheel mounting configurations.

Reliable and efficient vibroacoustic loads prediction is often critical in structural design, yet it remains a challenging task for many applications. Spacecraft structures are characterised by extensive use of composite materials, complex connections between components and various non-trivial geometrical features. Accurate treatment necessitates the construction of highly detailed numerical models, traditionally employing deterministic representations. Simultaneously, the broadband acoustic excitation due to the diffuse sound field experienced during launch requires modelling the fluid domain and solving the resulting elasto-acoustic interaction at = multiple frequencies.

To alleviate the computational demand implications for large problem sizes, substructuring and reduction techniques for the structural domain are commonplace, component mode synthesis (CMS) being a framework widely adopted in the aerospace industry. Nevertheless, despite ongoing research, the topic still presents a range of difficulties when a universal, robust method of accounting for model uncertainties is sought.

In this study, two CMS based approaches are proposed and evaluated. Firstly, the Craig-Bampton stochastic method (CBSM) is improved via a set of modifications enhancing its efficiency, and subsequently adapted for use in a vibroacoustic setting. Optimal perturbation levels and scope of validity of the technique are established against a probabilistic structural analysis (PSA) simulation for a spacecraft structure.

Secondly, a novel stochastic finite element method (FEM)
is presented. The underlying mathematical foundation is derived so that uncertainty can naturally be controlled at the subsystem level, in partitions of the corresponding condensed mass and stiffness matrices. This decomposition based approach ensures that realisations of the random matrices have key properties such as positive (semi)definiteness strictly preserved, guaranteeing complete robustness. The method is validated with a spacecraft test case, comparing its predictions against PSA, the improved CBSM and experimental data.
A coupling scheme with a hierarchical matrix accelerated boundary element method is formulated, resulting in the construction of a complete stochastic vibroacoustic solver.